Att_B_422-L1-STISSPEC-0002_Ver0.1_-_Signed.pdf

Effective Date: Oct. 3, 2019 422-L1-STISSPEC-0002

Expiration Date: TBD Version 0.1

Responsible Organization: SWFO-L1/Code 422 i

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422-L1-STISSPEC-0002

Version 0.1

Space Weather Follow On - Lagrange 1 (SWFO – L1)

Code 422

SWFO – L1

Supra Thermal Ion Sensor (STIS)

Requirements Specification (SPEC)

Goddard Space Flight Center

Greenbelt, Maryland

National Aeronautics and Space Administration https://goessp.ndc.nasa.gov/

SWFO STIS SPEC 422-L1-STISSPEC-0002

ii

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SWFO – L1 STIS Specification Requirements (SPEC)

Signature/Approval Page

Prepared by:

Michael Honaker Date

SWFO-L1 STIS Instrument Engineer

NASA/GSFC, Code 422

Reviewed by:

Ronald Hooker Date

SWFO-L1 Instrument Systems Manager

Approved by:

Gene Martin Date

SWFO-L1 Project Manager iii

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CM FOREWORD

This document is a Space Weather Follow On - Lagrange 1 (SWFO-L1) Project Configuration

Management (CM)-controlled document. Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the SWFO-L1 CM Office (CMO), along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

Questions or comments concerning this document should be addressed to:

NASA/Goddard Space Flight Center

SWFO-L1 Project Office, Code 422

Attention: Configuration Management Office

Greenbelt, Maryland 20771 iv

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Change History Log

REV

LEVEL DESCRIPTION OF CHANGE

APPROVED

BY

DATE

APPROVED

0.0 Initial release.

0.1 Updated section 4.1.5 to add “and the STIS instrument shall provide, as a minimum, one data collection in each 5 minute window.”

Updated section 4.2.8 to redefine the static magnetic field measurement requirement from “40 nT peak-to-peak in any axis 1 meter “ to “40 nT in any axis 1 meter” v

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Table of TBDs/TBRs/TBPs

Action Item

No.

Location Summary Individual/

Organization

Actionee vi

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Table of Contents

1.0 Introduction

1.1 General Information

1.2 Scope

2.0 Applicable Documents

3.0 Contract Description

3.1 STIS Description

3.2 Ground Support Equipment Description

4.0 Functional/Performance Requirements

4.1 STIS Performance Requirements

Maximum and Minimum Flux Flux Measurement Range

Accuracy Field of View (FOV) Refresh Rate and Latency Response Stability

4.2 Resource Allocations

Mass

Physical Envelope Operational Power Peak Operational Power

Survival Heater Power Telemetry

Transient Magnetic Field Static Magnetic Field

4.3 STIS Operating Modes

In-Flight Calibration

Mode Transitions Deterministic Power-on Configuration Safe Mode

4.3.4.1 Entry into Safe Mode

Fail-safe Recovery Mode Normal Operational Mode On-Orbit Operations

Station Keeping Activation

4.4 Power

Voltage Range Abnormal Voltages Power Transients Sudden Removal of Power

Over-Current Protection Primary Power Return Ground Turn-on Current Transients

Operational Current Transients vii

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4.5 Electrical Grounding

Primary Power DC Isolation Survival Power Isolation

Internally Generated Secondary to Primary DC Isolation Internally Generated Secondary Return Mechanical Contact Resistance Grounding Connector DC Resistance

4.6 Signal And Data Interfaces

Passive Analog Telemetry Data Signal Interface

4.6.2.1 Telemetry Source Packet Format

4.6.2.2 Command Source Packet Format

Clock Signal Interface Command and Housekeeping telemetry

Commands for Autonomous Functions Limits and Triggers

On-Board Processor Reset Microswitches

4.7 Flight Software

Flight Load Non-volatile Memory Software Updates

Software Table Updates Flexibility and Ease of Software Modification

Version Identifiers Warm Restart

Memory Tests Memory Dump

Fault Detection and Correction Control Health and Safety Monitoring

5.0 Physical Requirements

5.1 Interface and Design Units and Exculsions

5.2 Mass Properties Accuracy

STIS Instrument Mass Accuracy Center of Mass Location Center of Mass Accuracy Determination of Moments and Products of Inertia

5.3 Mounting

Surface Flatness Method

5.4 Coordinate System and Alignment

6.0 Environmental Requirements

6.1 Mechanical Factors of Safety

6.2 Quasi-Static Acceleration

6.3 Frequency Requirement

Fundamental Launch Frequencies viii

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6.4 Vibration

Sinusoidal Vibration Random Vibration

6.5 Shock

6.6 Acoustics

6.7 Transportation

Transportation Cleanliness

6.8 Pressure

Operating Pressure Range Maximum Depressurization Rate Launch Vehicle (LV) Environmental Control System (ECS) Impingement

6.9 On-Orbit Dynamic Environment

6.10 Thermal Requirements

Flight Interface Design Temperature Limits

6.11 Charged Particle Radiation Requirements

Definitions Non-Destructive Events (SEUs, SETs, SEFIs, SHEs, and MBUs)

6.11.2.1 Single-Event Effect Rate Calculations

Destructive Events

6.11.3.1 Single Event Latchups (SELs)

6.11.3.2 Single Event Gate Rupture (SEGRs), Single Event Burnout (SEB)

Charging Environment

Total Ionizing Dose Displacement Damage Dose

7.0 Contamination control

7.1 STIS Surface Cleanliness Requirements

7.2 Contamination Generation

Particulate Generation

Molecular Contamination Generation

7.2.2.1 Molecular Material Restrictions

7.3 Vacuum Bakeouts

7.4 Contamination Analyses

7.5 Venting Requirements

7.6 Instrument Purging (including during integrated Observatory activities)

7.7 Cleanability and Protection

7.8 Electrostatic Cleanliness

Conductive Surface Ground Path Conductive Surface Resistivity

Closeout of Gaps and Apertures Exposed Harness Specific Requirements

Thermal Blankets

8.0 Design & Construction Requirements

8.4 Electrical

Test Sensors MGSE Grounding Connector Specifications ix

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8.4.3.1 Contact Derating

8.4.3.2 Redundant Contact Derating

8.4.3.3 Signal Segregation

8.4.3.4 Test and Flight Signal Isolation

Test Interfaces

8.4.4.1 Facility-Induced Noise

8.4.4.2 Facility-Induced ESD GSE Malfunction

8.4.4.3 Facility-Induced GSE Malfunction

Mitigation of Internal Charging

8.4.5.1 Mitigation Strategies for Internal Charging

8.4.5.2 Floating Conductors

8.4.5.3 Dielectric Structures

8.5 Safety

8.6 Electromagnetic Compatibility

Conducted Emissions

8.6.1.1 Applicability of Conducted Emissions

8.6.1.2 CE101 – Differential Mode Current Emission Limits

8.6.1.3 CE102 – Differential Mode Current Emission Limits

Common Mode Current Emissions Limits

Conducted Susceptibility

8.6.3.1 Applicability of Conducted Susceptibility

8.6.3.2 CS101 – Power Leads, 30 Hz to 150 kHz Limit

8.6.3.3 CS114 – Power Leads, 150 kHz to 50 MHz Limit

8.6.3.4 CS114 – Power and Signal Cables, Common Mode

Radiated Emissions

8.6.4.1 RE102 - Electric Field Emissions

Radiated Susceptibility

8.6.5.1 RS103 - Electric Field

8.7 Identification and Marking

8.8 Workmanship

Connectors

8.9 Reliability and Mission Lifetime

Mission Life

Operating Time Trouble-Free Time

8.10 Ground Handling

Ground Support Equipment (GSE) Design Lifting Hardware

Manual Lifting Hardware GSE Cleanliness

GSE Bakeout Test Harness

8.11 Interface Documentation

Mechanical Interface Electrical Interface Data Interface x

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9.0 Mechanical Design Requirements

9.1 Structural Requirements

Component Fatigue

Fracture Control Requirements

9.2 Fastening Systems

Factors of Safety Supplemental Factor Ultimate Design Loads

Yield Design Loads Design Separation Load Fastener Locking and Retention

9.2.6.1 Thread Engagement

9.2.6.2 Locking Feature Verification

9.2.6.3 Locking Feature Installation

9.2.6.4 Snap Ring & Cotter Pin Use Limitation

9.2.6.5 Snap Ring & Cotter Pin Use

9.2.6.6 Liquid Locking Compounds

Fastened Joints Criteria

10.0 Logistics

10.4 I&T Deliverables

10.5 Ground Support Equipment

Electrical System Test Equipment

STIS Emulator (STISE)

10.6 Transportation Equipment

Shipping Container

11.0 Verification Requirements

11.4 Verification Methods

Inspection

Analysis Test

11.5 Inspection Requirements

Visual Inspection Physical Measurement

Documentation Search

11.6 Analysis Requirements

11.7 Test Requirements

Definitions Test Factors

Test Tolerances Test Restrictions

11.8 Required Tests

Performance Tests Mass Properties Measurement Static Loads/Strength Test Sine Sweep Survey Sine Vibration xi

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Random Vibration

Acoustic Test Shock

Thermal Vacuum Bake-out Thermal Vacuum Test Magnetics Test Harness Tests EMI/EMC Tests

Appendix A Abbreviations and Acronyms xii

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List of Figures

Figure Page

Figure 6-1 Sine Vibration Environment [TBR]

Figure 6-2 STIS Acceptance [TBR] Random Vibration Environment Figure 6-3 SWFO-L1 S/C Acceptance [TBR] Shock Envelope Figure 6-4 SWFO-L1 S/C Protoflight [TBR] Acoustic Envelope Figure 6-5 Total ionizing dose-depth curve for a 5-year mission at L1 Figure 6-6 Silicon displacement damage dose-depth curve for a 5-year mission at L1

Figure 6-7 Gallium Arsenide displacement damage dose-depth curve for a 5-year mission at L1 ..

Figure 8-1 CE101/CE102 Differential Mode Current Emission Limits Figure 8-2 Common Mode Current Emission Limits Figure 8-3 Power Lead Conducted Susceptibility (CS101) Voltage Limit Figure 8-4 Unit Level RE102 Radiated Electric Field Emission Limits [TBR]

Figure 11-1 Thermal Vacuum Profile

List of Tables

Table Page

Table 2-1 Applicable Documents

Table 6-1 Factors of Safety Table 6-2 STIS Design Limit Loads [TBR]

Table 6-3 Qualification [TBR] Level Shock Response Spectrum Table 6-4 Transportation Loads Table 6-5 Temperature Limits at Instrument Mounting Interface Table 6-6 Dose requirements as a function of shielding for TID in silicon, DDD in silicon and

DDD in gallium arsenide Table 7-1 STIS Lifetime Contamination Requirements Table 7-2 Thermal Blanket Area vs. Grounding Tabs Table 8-1 Emission and Susceptibility Requirements Table 8-2 RS103 Radiated Susceptibility Levels

Table 11-1 Test Factors and Durations

Table 11-2 Test Tolerances

Table 11-3 Thermal Vacuum Test Parameters

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1.0 INTRODUCTION

1.1 GENERAL INFORMATION

1.2 SCOPE

This specification describes the performance and electrical, mechanical, environmental, and verification testing requirements for a space-qualified SupraThermal Ion Sensor (STIS) for the

Space Weather Follow-On at Lagrange 1 (L1) Mission.

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2.0 APPLICABLE DOCUMENTS

The following documents and drawings in effect on the day this specification was signed shall apply to the fabrication and to the electrical, mechanical, and environmental requirements of the

STIS to the extent specified herein. In the event of conflict between this specification and any referenced document, this specification will govern, with the exception of the SWFO-L1 STIS

Statement of Work (422-L1-STISSOW-0001), in which case the Statement of Work takes precedence.

The following is a list of the applicable specifications and publications.

Table 2-1 Applicable Documents

Document Number Title

422-L1-STISSOW-0001 SWFO-L1 STIS Statement of Work

422-L1-STISCDRL-0003 SWFO-L1 STIS Deliverable Items List and Schedule (DILS)

ANSI/TIA/EIA-422-B Electrical Characteristics of Balanced Voltage Digital Interface Circuits

ANSI/TIA/EIA-644-A-

Electrical Characteristics of Low Voltage Differential Signaling (LVDS)

Interface Circuits

NFPA 70 National Fire Protection Association National Electric Code

NASA-STD-5001B Structural Design And Test Factors Of Safety For Spaceflight Hardware

NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety Requirements

NASA-STD-6016 Standard Materials and Processes Requirements for Spacecraft

NASA-HDBK-7005 Dynamic Environment Criteria

NASA-STD-7001 Payload Vibroacoustic Test Criteria

IEST-STD-CC-1246E Product Cleanliness Levels And Contamination Control Program

ASTM E-595-07 Standard Test Method for Total Mass Loss and Collected Volatile

Condensable Materials from Outgassing in a Vacuum Environment

MIL-DTL-5541 Chemical Conversion Coatings on Aluminum and Aluminum Alloys

MIL-A-8625F Anodic Coatings for Aluminum and Aluminum Alloys

EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating

MIL-STD-461G Military Standard, Electromagnetic Emission And Susceptibility

Requirements For The Control Of Electromagnetic Interference (EMI)

GSFC-STD-7000A General Environmental Verification Standard (GEVS)

NASA-STD-5019A Fracture Control Requirements for Spaceflight Hardware

NASA-STD-5020 Requirements for Threaded Fastening Systems in Spaceflight Hardware

NASA-STD-5017A Design and Development Requirements for Mechanisms

NASA-HDBK-4002A Mitigating In-Space Charging Effects—A Guideline

FAA AC 20-71 Federal Aviation Administration Advisory Circular (AC) 20-71, “Dual

Locking Devices on Fasteners".

NASM 33540 Safety Wiring, Safety Cabling, Cotter Pinning, General Practices for

SAE AS567 Safety Cable, Safety Wire, Key Washers, and Cotter Pins for Propulsion

Systems, General Practices for Use of

541-WI-5330.1.41 Fastener Locking Using Arathane 5753

MSFC-STD-3029A Guidelines for the Selection of Metallic Materials for Stress Corrosion

Cracking Resistance in Sodium Chloride Environments

MIL-STD-462, Notice 2 Electromagnetic Interference Characteristics, Measurement of, 1 May 1970

ISO 14644-1:2015 Part 1: Classification of air cleanliness by particle concentration

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Document Number Title

ISO 14644-3: 2005 Cleanrooms and associated controlled environments -- Part 3: Test methods

CCSDS 133.0-B-1 Space Packet Protocol, Blue Book, Issue 1, September 2003

CCSDS 301.0-B-4 - Time Code Formats. Blue Book. Issue 3, January 2002

3.0 CONTRACT DESCRIPTION

3.1 STIS DESCRIPTION

The SupraThermal Ion Sensor (STIS) is an ion spectrometer device that measures ions across a broad range of energies to characterize solar ejecta including Coronal Mass Ejections (CMEs), co-rotating interaction regions (CIRs) and interplanetary shocks. In particular, STIS is a low energy charged particle detector that is capable of measuring the ion flux population as a function of energy. This ion population with energies higher than that of the bulk plasma are produced by local solar acceleration as well as acceleration from a CME shock front. Analysis of these spectra can aid in estimating the arrival time and strength of CMEs shocks.

The STIS contract includes the delivery of one (1) flight model (FM) instrument and one (1)

Engineering Development Unit (EDU) [TBR], flight harness(es) between instrument boxes, if applicable, and enough spares and built-up sub-assemblies for one additional FM [TBR].

3.2 GROUND SUPPORT EQUIPMENT DESCRIPTION

The STIS contract includes the delivery of (2) sets [TBR] of the Electrical System Test

Equipment (ESTE), (2) [TBR] STIS Emulators (STISEs), and (1) One Ground Processing

Development System (GPDS). Additionally, Mechanical Ground Support Equipment (MGSE) such as lifting fixtures/handles, shipping containers, purge carts, drill templates, test fixtures, non-flight protective covers, etc. are included in the delivery. Electrical Ground Support

Equipment includes, but is not limited to, test cables and break-out boxes as required.

Ground Support Equipment is necessary to operate the instrument during spacecraft testing.

Supplying this equipment is considered part of the delivery for the instrument. More information is provided in the contract deliverable requirement list (CDRL).

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4.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the STIS.

The requirements in this Specification pertain to the STIS ‘system’, which may include all instrument hardware, software, and ground processing algorithms. The STIS contractor is not responsible for the operational ground system, but certain specifications will require ground processing after collection but before data distribution for which the contractor is responsible for defining.

4.1 STIS PERFORMANCE REQUIREMENTS

Maximum and Minimum Flux

The STIS requirements below utilize the terms Minimum Flux and Maximum Flux. These terms are defined as follows:

Flux ([cm2 sec sr KeV]-1):

Minimum Flux: 2.48x102 * E(KeV)-2.3

Maximum Flux: 1.01x107 * E(KeV)-1.6

Flux Measurement Range

The STIS shall provide ion differential flux measurements in the range of 15KeV - 2,000 KeV.

Discrimination of flux constituents or species is not required.

The STIS shall determine the differential ion fluxes in sufficient evenly spaced logarithmic energy bands to meet the accuracy requirement defined in section 4.1.3.

Accuracy

The STIS instrument shall be capable of measuring differential fluxes between the minimum and maximum flux with 20% accuracy above 22 keV [TBR] and 45% accuracy between 15keV and

22keV [TBR] accounting for all error sources. The instruments shall provide a number of energy bands sufficient to support the differential ion flux accuracy requirements at all energies.

Error sources include but are not limited to noise, out-of-band response, energy band edge uncertainty, poisson statistics, and facility calibration error.

Field of View (FOV)

The STIS FOV shall be at least 80 degrees in the ecliptic by 60 degrees centered 50 degrees in the ecliptic off of the sun-earth line in the “ahead” direction (the ahead direction is the direction the earth travels in the ecliptic). In the Spacecraft Reference Frame (SRF) coordinates (see section 5.4), the unit vector for the centroid of the FOV of the instrument shall be (0.643, 0, 0.766).

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The spacecraft will provide an unobstructed field of view.

Refresh Rate and Latency

The SWFO-L1 mission will make STIS Solar Wind data available to SWPC forecasters within 5 minutes of the completion of each data acquisition and the STIS instrument shall provide, as a minimum, one data collection in each 5 minute window.

The processing of a completed on-orbit data collection by the STIS instrument and the transmission to the spacecraft shall contribute ≤ 2 [TBR] seconds to the overall latency of the solar wind data product.

The STIS ground processing algorithms contribution to data latency of the STIS Level 1b algorithm output shall be ≤ 236 [TBR] seconds.

Response Stability

The STIS instrument shall not saturate at fluxes up to three time the maximum flux for a given energy range. Accuracy is allowed to degrade when measuring fluxes above the maximum flux value defined, but the instrument shall respond with quantifiable measurements up to 3 times the maximum flux to ensure that such high flux signals are real and not the result of a malfunction.

4.2 RESOURCE ALLOCATIONS

Mass

The STIS shall have a mass of less than or equal to 5 kg [TBR].

Physical Envelope

The STIS volume (length*width*height), including mounts, thermal blankets and connectors, for both stowed and operational configurations shall have dimensions that do not exceed a total volume of 7500 cm3 [TBR]. These dimensions pertains to both static and dynamic envelopes of the instrument.

The STIS maximum dimension shall be less than 25 cm [TBR].

Operational Power

The STIS shall have a nominal operational power of ≤4W.

Peak Operational Power

The STIS shall have a peak operational power of ≤ 5 W [TBR] over a maximum duration of 60 seconds.

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Survival Heater Power

The STIS shall have a survival heater power consumption of ≤ 4W [TBR].

Telemetry

The STIS instrument science and engineering data rate, including all overhead associated with

Consultative Committee for Space Data Systems (CCSDS) packetization by the instrument at the spacecraft interface, shall not exceed 1 [TBR] Kbits per second (Kbps) maximum.

Transient Magnetic Field

The STIS instrument shall limit varying magnetic field less than 5 Hz [TBR] to less than 4 nT

[TBR] zero-to-peak in any axis 1 meter from any face of the unit for any operating mode and changing configurations.

Static Magnetic Field

The STIS units shall limit the static magnetic field to less than 40 nT in any axis 1 meter from any face of the unit for any operating mode [TBR].

4.3 STIS OPERATING MODES

The STIS current operating mode shall be identified by a flag in its telemetry stream.

In-Flight Calibration

The STIS shall provide an in-flight calibration mode that provides a test input to support calibration and testing both on the ground and in space.

The STIS instrument shall still be capable of providing normal science data while operating in this mode.

The STIS in-flight calibration shall be both self-terminating and able to be terminated by a ground command.

Mode Transitions

The STIS instrument shall transition from the current mode to any other mode without causing damage to itself.

Deterministic Power-on Configuration

The STIS On-Board Processor shall initialize upon power-up into a predetermined configuration.

This predetermined configuration shall not enable any high voltage power supplies. High voltage power supplies shall be enabled by ground command only.

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Safe Mode

The STIS instrument shall provide a Safe Mode. Safe mode is an operating mode of the instrument during which all non-essential subsystems are shut down and only essential functions are active. Safe mode represents a configuration in which the instrument is thermally, mechanically, and optically “safe” without receiving commands from the Spacecraft.

In safe mode any STIS high-voltage power supplies shall be disabled. High voltage power supplies shall be re-enabled by ground command only.

The STIS instrument shall be capable of remaining in a safe configuration for at least 120 hours without ground intervention.

4.3.4.1 Entry into Safe Mode

The STIS instrument shall enter Safe Mode upon detection of internal faults that are capable of causing damage to the instrument.

The SWFO-L1 spacecraft will monitor up to four analog instrument health and safety parameters defined by the STIS contractor and command the STIS into safe mode when any of those key values have been exceeded.

The SWFO-L1 spacecraft will monitor up to 12 [TBR] digital telemetry points that are part of the normal instrument generated data stream for instrument health and safety parameters defined by the STIS contractor and command the STIS into safe mode when any of those key values have been exceeded.

Fail-safe Recovery Mode

The STIS instrument shall provide a failsafe recovery mode dependent on a minimal hardware configuration capable of accepting and processing a minimal command subset sufficient to load and dump memory.

Normal Operational Mode

In normal operational mode, the STIS instrument shall be in a fully functional configuration in which designed measurements are made in accordance with the performance requirements listed in section 4.1 of this document and are available to be sent to the spacecraft for downlink.

On-Orbit Operations

The STIS instrument shall operate normally, within specification, while flying aboard a 3-axis stabilized spacecraft with orbital limit constraints as stated in this specification.

Station Keeping

The STIS instrument shall continuously operate during all spacecraft maneuvers.

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The STIS instrument may meet performance requirements during these maneuvers but shall meet them again after each maneuver is completed within 300 [TBR] secs.

Activation

The STIS instrument shall require no active commanding prior to fourteen (14) days [TBR] after launch.

4.4 POWER

Voltage Range

The STIS shall operate over the bus voltage range of +27 to +35 VDC at the primary power inputs during all normal mission phases and for all expected load conditions (except when turned off).

Abnormal Voltages

The STIS shall survive without performance degradation after indefinite exposure to an anomalous voltage range of 0 to +40 VDC.

Power Transients

The STIS shall meet its performance requirements in the presence of transients specified in MIL-

STD-461G, Figure CS115-1 and Figure CS116-2.

Sudden Removal of Power

The STIS shall meet its performance requirements without degradation after exposure to an abrupt, unannounced removal of power.

Over-Current Protection

The STIS shall not use non-resetting over-current protection (i.e., fuses) internal to the unit.

Primary Power Return Ground

STIS shall provide a dedicated Primary Power return in the same connector as the primary power.

Turn-on Current Transients

The STIS shall be capable of receiving limit turn-on input voltage (0 – 35V) with a rise time of

500us or longer.

The STIS shall limit its instrument operational power turn-on current ramp rate to less than

2A/microsecond.

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The STIS instrument shall limit the instrument operational power turn-on current in-rush transient(s) to 1.0A [TBR] for a period less than 20 milliseconds [TBR].

Following initial in-rush, the STIS instrument shall limit the instrument operational power current transient(s) to 0.4A peak to peak [TBR].

Note: The spacecraft bus will not drop below 27V during the STIS instrument turn on.

Operational Current Transients

The STIS instrument shall limit any change in operational power current at any time (including initial power turn-on) to less than 0.2A/µs [TBR].

The STIS instrument shall limit the maximum delta change in operational current at any time to

0.1 A [TBR].

4.5 ELECTRICAL GROUNDING

Primary Power DC Isolation

The STIS primary power interfaces, primary power and primary power returns, shall be isolated from the unit chassis by a DC resistance of greater than or equal to 10 Megaohms.

Survival Power Isolation

The STIS survival heater power interfaces shall be isolated from the unit chassis by a DC resistance of greater than or equal to 10 Megaohms.

Internally Generated Secondary to Primary DC Isolation

Secondary power (or signal) inputs shall be isolated from primary power by a DC resistance of greater than 10 Megaohms.

Internally Generated Secondary Return

The STIS instrument shall reference its secondary returns (power and signal grounds) to the unit chassis ground by connecting them at one or more places.

Mechanical Contact Resistance

The DC resistance of the mechanical contact between two conductive mating surfaces (internal to the unit) shall be less than or equal to 2.5 mΩ DC resistance.

Grounding

The DC resistance between the STIS chassis and the observatory chassis shall be ≤ 2.5 milliohms.

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The STIS shall provide a ground lug for a grounding strap to be attached from the STIS chassis for connection to the spacecraft conductive structure [TBR].

The grounding lug location on the STIS instrument chassis or the tie points in contact with the ground strap shall have a minimum contact area of 80 mm². The ground lug contact area must remain free of any material finish that may affect the reliability of the ground connection and will be shown in the MICD.

Connector DC Resistance

STIS connectors shall be electrically connected to chassis with a DC resistance ≤ 2.5 mΩ.

4.6 SIGNAL AND DATA INTERFACES

Passive Analog Telemetry

The STIS shall utilize no more than 6 [TBR] analog signals to monitor critical temperature points when the STIS is powered off.

Data Signal Interface

The STIS Instrument shall interface for data transfer to or from the spacecraft by either RS-422 or Low Voltage Differential Signal ECS-E50-12A (LVDS) data bus complying with (TBP), or

ANSI/TIA/EIA-422-B Electrical Characteristics of Balanced Voltage Differential Interface

Circuits, or ANSI/TIA/EIA-644-A-2001 Electrical Characteristics of Low Voltage Differential

Signaling (LVDS) Interface Circuits, as applicable.

4.6.2.1 Telemetry Source Packet Format

The STIS shall transfer all data to the spacecraft using the CCSDS 133.0-B-1 Section 4.1

Protocol Data Unit definition shown in the Telemetry Source Packet Definition Figure.

TELEMETRY SOURCE PACKET DEFINITION

SEC.

HDR

FLAG

13 - 8K octets

3 bits 1 bit 1 bit 11 bits 2 bits 14 bits 16 bits 72 bits

DATA

VARIABLE

TIME CODE PACKET

LENGTH

SECONDARY

HEADER

PACKET

SEQUENCE

COUNT

SEQUENCE

FLAGS

PRIMARY HEADER

PACKET IDENTIFICATION PACKET SEQUENCE CONTROL

VERSION

NUMBER

TYPE APPLICATION

PROCESS ID

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The STIS telemetry Source packets shall be variable length with a maximum data zone of 8192 octets including Secondary Header.

The STIS shall set the telemetry source packet Secondary Header Flag to the value 1.

The STIS shall set the Telemetry Source Packet Sequence Flags to the value of 11.

Note: Segmentation services are not permitted.

The STIS shall set the Telemetry Source Packet Time Code per CCSDS 301.B-4 Time Code

Formats, Day Segmented format in the Time Code Format Figure.

Note: The P-Field is implied and not included in the actual time message.

The spacecraft on-board reference time will be International Atomic Time (TAI) [TBR]

4.6.2.2 Command Source Packet Format

The STIS shall receive all data from the spacecraft formatted per CCSDS 133.0-B-1 Section 4.1

Protocol Data Unit definition shown in the Command Source Packet Definition Figure.

COMMAND SOURCE PACKET FIGURE

The STIS command Source packets shall be variable length with a maximum data zone of 8192 octets.

The STIS shall receive Command Source Packet with Secondary Header Flag set to the value 0.

The STIS shall receive Command Source Packet with the Sequence Flags set to the value of 11.

Note: Segmentation services are not permitted.

SEC.

HDR

FLAG

13 - 8K octets

3 bits 1 bit 1 bit 11 bits 2 bits 14 bits 16 bits

DATA

VARIABLE

PACKET

LENGTH

PACKET

SEQUENCE

COUNT

SEQUENCE

FLAGS

PRIMARY HEADER

PACKET IDENTIFICATION PACKET SEQUENCE CONTROL

VERSION

NUMBER

TYPE APPLICATION

PROCESS ID

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Clock Signal Interface

The STIS shall receive a Pulse per Second time code sequence from spacecraft clock line by either RS-422 or Low Voltage Differential Signal (LVDS) data bus complying with

ANSI/TIA/EIA-422-B Electrical Characteristics of Balanced Voltage Differential Interface

Circuits, or ANSI/TIA/EIA-644-A-2001 Electrical Characteristics of Low Voltage Differential

Signaling (LVDS) Interface Circuits, as applicable.

The spacecraft will provide STIS a 1 PPS time code sequence accurate to +/- 1 msec [TBR] relative to TAI.

The STIS shall receive from the spacecraft a time code message on the data line as defined in

Spacecraft Time Message Packet Figure. The time code message is the time applicable to receipt of the 1 PPS.

Command and Housekeeping telemetry

The STIS instrument shall provide command and housekeeping telemetry functions in all powered modes.

Commands for Autonomous Functions

The STIS shall execute commands to individually enable and disable each autonomous function.

Limits and Triggers

The STIS autonomous limits and triggers shall be changeable by command.

On-Board Processor Reset

The STIS On-Board Processor shall be reset by command.

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An autonomous power-on reset occurrence shall be unambiguously identifiable via telemetry.

Note: This does not imply real-time telemetry as the reset is occurring.

Microswitches

Micro-switches shall only be used for information only and not use to initiate on-board autonomous activity or as an on-board interlock.

4.7 FLIGHT SOFTWARE

Flight Load Non-volatile Memory

The STIS flight software image shall be contained in its entirety in non-volatile memory at launch.

Software Updates

The flight software modules shall be reprogrammable.

Activation of uploaded modified software shall not require an upload of the entire flight software image [TBR].

Software Table Updates

Instrument configuration data (e.g. Table Loads or Configuration Parameters) shall be reprogrammable during integration and test phases and on-orbit without computer restart [TBR].

Modified instrument configuration data (e.g. Table Loads or Configuration Parameters) shall be committed to operational use by ground command.

Flexibility and Ease of Software Modification

The STIS flight software shall be deterministic in terms of scheduling and prioritization of critical processing tasks to ensure their timely completion.

Instrument configuration data (e.g. Table Loads or Configuration Parameters) shall be referenced such that data can be loaded and dumped by the ground without reference to memory address.

The definition of instrument commands within the ground database (excluding memory diagnostic/load commands) shall not be dependent on physical memory addresses within the flight software [TBR].

Version Identifiers

The STIS software and firmware versions shall each have an internal identifier (embedded in the executive program) that can be included in the instrument engineering data.

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This software identifier shall be configuration management controlled.

Warm Restart

The STIS flight software shall provide a restart by command with preservation of instrument configuration data and memory tables.

Memory Tests

The STIS flight software shall provide a mechanism to verify the contents of all memory areas.

Memory Dump

The STIS flight software, and associated on-board computer hardware, shall provide the capability to dump any memory location.

The flight software memory dump capability shall not disturb normal operations and instrument data processing [TBR].

Fault Detection and Correction Control

If applicable, STIS shall provide the capability to enable and disable any internal Fault Detection and Correction (FDC) features.

Health and Safety Monitoring

The STIS flight software shall provide health and safety monitoring, including memory checksum and watchdog timer, during integration and test phases and on-orbit.

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5.0 PHYSICAL REQUIREMENTS

5.1 INTERFACE AND DESIGN UNITS AND EXCULSIONS

The contractor shall use metric units to design the instrument and for developing interfaces with the spacecraft including any drawings, documents, models, except for the following cases:

Heritage Component or unit: Components or units that has been previously qualified, or of similar design heritage, may be specified in English units where use of metric equivalents would lead to additional cost to the program.

Fasteners: Although bolt patterns will be defined using metric dimensioning, use of

English fasteners (with hole dimensioning and tolerancing) is permitted.

Angular Measurement: Angular measurement may be expressed in degree of arc or in an appropriate subdivision of degree of arc such as second of arc (arc-sec) when advantageous to application.

5.2 MASS PROPERTIES ACCURACY

STIS Instrument Mass Accuracy

The mass of the STIS shall be measured to within ±0.1 kg.

Center of Mass Location

The contractor shall define the center of mass in the Mechanical ICD.

Center of Mass Accuracy

The center of mass of the STIS shall be determined to within ±5 mm [TBR] relative to a reference that will be defined in the MCID.

Determination of Moments and Products of Inertia

The Contractor shall determine by analysis the launch and on-orbit moments and products of inertia to an accuracy of ±5.0 percent [TBR] of the maximum principal moment of inertia, referenced to the coordinate axes with an origin at the center of gravity.

5.3 MOUNTING

Surface Flatness

Mounting interface flatness, and co-planarity requirements for the STIS side of the interface

(including brackets, if any, and shims) shall be as defined in the MICD.

Method

The method by which the STIS is mounted to the spacecraft will be defined in the Spacecraft to

STIS Interface Control Document (ICD).

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5.4 COORDINATE SYSTEM AND ALIGNMENT

The SWFO-L1 spacecraft on-orbit coordinate system uses a right-hand orthogonal, body fixed

XYZ coordinate system (Spacecraft Reference Frame - SRF) as follows: the positive XSRF-axis points toward the Sun, the positive YSRF-axis points towards the north ecliptic pole and the ZSRF axis is obtained from the right hand rule. The roll, pitch and yaw rotations are defined about the

X, Z, and Y axes, respectively. The origin of coordinate system is at the center of the ESPA ring attachment plane.

The spacecraft will align the sensor axes to within +/-2 [TBR] degrees of the direction specified in section 4.1.4.

The STIS shall have clear fiducial marks to enable alignment during integration with spacecraft.

The instrument unit mounting frame is an orthogonal reference frame defined by the locations of the spacecraft side of the instrument unit mounting points. A rigorous definition of this frame will be documented in the ICD.

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6.0 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements for the STIS Instrument are specified in this section.

The STIS shall meet its performance requirements in section 4.0 after exposure to the environments specified in this section.

6.1 MECHANICAL FACTORS OF SAFETY

The STIS flight hardware shall demonstrate positive Margins of Safety under limit loads for all yield and ultimate failures using the Factors of Safety (FS) defined in Table 6-1. Margin of

Safety (MS) is defined as follows:

MS = (Allowable Stress (or Load) / (Applied Limit Stress (or Load) x FS)) -1

Table 6-1 Factors of Safety

Primary and secondary structure comprised of composite materials, Beryllium, bonded joints and/or bonded inserts shall be proof tested to 1.25 x Limit Load; qualification by analysis only is not acceptable. Actual flight hardware testing is preferred, but testing of representative sets of hardware with a similar qualification argument can be used if approved by the NASA/GSFC

COR.

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6.2 QUASI-STATIC ACCELERATION

Quasi-static acceleration represents the combination of steady-state accelerations and the low frequency mechanically transmitted dynamic accelerations that occur during launch.

The STIS shall be designed to withstand the quasi-static design limit loads defined in the mass-acceleration curve (MAC) shown in Table 6-2 without damage or degradation of performance and are to be applied one axis at a time. The design loads shown below will be updated based on the results of coupled loads analysis.

Linear interpolation should be used between breakpoints to determine the appropriate limit load as a function of STIS weight. Note that these design limit loads are intended to cover only the low frequency launch environment and must be used in conjunction with the random vibration environments to assess structural margins.

Table 6-2 STIS Design Limit Loads [TBR]

MAC Breakpoints

Mass (kg) GSFC (g) Moog* RSS (g)

1.0 51.3 125.7

5.0 42.4 60.8

10.0 36.6 44.5

20.0 30.3 32.5

40.0 24.2 23.8

60.0 20.9 19.8

80.0 18.7 17.4

100.0 17.2 15.7

120.0 15.9 14.5

140.0 15.0 13.5

160.0 14.2 12.7

181.0 13.5 12.0

200 12.9 11.5

300 10.8 9.6

400 9.5 8.4

Notes:

1. Moog MAC values are taken from “ESPA User’s Guide, The EELV Secondary Payload

Adapter, November 2018

2. Numbers in red are extrapolated from the Moog MAC

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6.3 FREQUENCY REQUIREMENT

Fundamental Launch Frequencies

The STIS shall have a fundamental frequency greater than 100 Hz (TBR) when hard mounted at its spacecraft interface. Any unit, which fails to meet the specified fundamental frequency, must supply a finite element model, correlated to modal survey test results up to 50 Hz, to be used in coupled loads analyses. Requirements for the submitted finite element model are shown in the

Magnetometer SOW and discussed in Section 11.8.4 of this document.

6.4 VIBRATION

Sinusoidal Vibration

The STIS shall undergo qualification, protoflight or acceptance (level depends on qualification status or qualification approach for the unit) sine vibration testing in all three axes at the levels shown in Figure 6-1. Instruments with a first mode greater than 150 Hz can be exempted from sine vibration testing upon approval by the NASA/GSFC COR. A generic sine vibration specification is provided for acceptance levels applied at the SWFO-L1 to STIS interface is shown (Note: in Figure 6-1, axial refers to the launch vehicle thrust axis and lateral to either of the other two orthogonal directions). See Section 11.7.1 for definitions of Protoflight, Qualification, and Acceptance.

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Figure 6-1 Sine Vibration Environment [TBR]

Levels may be notched to not exceed 1.25 times the design limit load outlined in section 6.2.

Peak levels at the low end of the frequency range (5 – 20 Hz typically) may be ramped up as needed to accommodate shaker table displacement limitations.

Random Vibration

The STIS shall demonstrate its ability to meet its performance requirements after being subjected to the random vibration environment in Figure 6-2 for units weighing 22.7 kg (50 lb.)

or less, applied at the Spacecraft to STIS interface.

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Figure 6-2 STIS Acceptance [TBR] Random Vibration Environment

The random environment will be updated by NASA once more information is available at the mission level. Note for lightweight STIS, the highest design loads may be from this random vibration environment.

The contractor shall provide random vibration analysis along with static loads analysis. Please see NASA-HDBK-7005 and NASA-STD-7001 for more information.

During the random vibration test, the test input level will be reduced (notched) at critical frequencies, if required, to limit the random vibration loads and/or acceleration responses to 3 dB above design limit levels.

Notching shall be limited to -12 dB of the original input and to a bandwidth of less than 100 Hz to limit the random vibration responses to 3dB above design limit levels. Notching beyond these limits will require NASA/GSFC COR approval.

6.5 SHOCK

The STIS shall be designed to meet its performance requirements after being subjected to the shock environment in Table 6-3, applied at the STIS interface to the SWFO-L1 spacecraft structure.

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Table 6-3 Qualification [TBR] Level Shock Response Spectrum

Freq (Hz) SRS (G)

100 100

800 1700

8000 1700

10000 2200

A shock susceptibility and attenuation assessment (analysis) shall be performed on all Spacecraft units or instruments. If the flight shock environment as shown on a Shock Response Spectra

(SRS) plot (Q=10) is enveloped by the curve shown below, then the shock environment can be considered benign and there is low risk in deferring the shock test to the Observatory level and the STIS contractor can request a waiver to not perform the shock test at the instrument level.

Analysis supporting this conclusion (i.e. deferral of shock testing) shall be provided to the

NASA/GSFC COR for review and approval with the waiver request.

Figure 6-3 SWFO-L1 S/C Acceptance [TBR] Shock Envelope

Any unit determined to be susceptible to the shock environment (e.g., where shock levels are expected to be above the shown curve) shall have shock testing performed at the unit level

(preferably on a qualification unit).

Unit self-induced shock testing shall be accomplished by two actuations at the unit level for each self-induced shock source (in order to account for the scatter associated with the actuation of the device) for the first flight unit, and a single actuation on subsequent units.

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6.6 ACOUSTICS

The STIS shall be designed to withstand, without any damage or degradation of performance, the equivalent Protoflight [TBR] levels shown in Figure 6-4. The STIS contractor shall use envelope option 1 for the SWFO-L1 design, analysis and testing if required. Note: Envelop option 1 and 2 in Figure 6-4 are the same above 800 Hz.

Figure 6-4 SWFO-L1 S/C Protoflight [TBR] Acoustic Envelope

An acoustic test shall be performed unless an assessment of the unit indicates that it is not susceptible to the expected acoustic environment, or responses are enveloped by random vibration testing, or that testing at higher levels of assembly provides sufficient exposure at an acceptable level of risk to the program as determined by the NASA/GSFC COR.

Unit acoustic tests shall be to Protoflight levels (Acceptance + 3 dB) with a duration of 1 minute.

6.7 TRANSPORTATION

In addition to the launch loads shown above, the STIS shall also be designed to withstand the maximum transportation loads shown in Table 6-4 without damage or degradation of performance.

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Table 6-4 Transportation Loads

Transportation Cleanliness

Materials and enclosure used for transportation and storage shall not generate molecular or particle contaminants or degrade the surface cleanliness of the item or adjacent items to level below those required in section 7.

6.8 PRESSURE

Operating Pressure Range

The STIS shall be designed to be capable of aliveness testing over a pressure range of 1.08 x 105

N/m2 (813 Torr) to 1.3 x 10-12 N/m2 (1 x 10-14 Torr).

The STIS shall meet all performance requirements while under vacuum 1.3 x 10-12 N/m2 (1 x 10-

14 Torr).

Maximum Depressurization Rate

The STIS shall be designed to meet all performance requirements after exposure to a maximum depressurization rate of -50 mbar/sec (-0.72 psi/sec) experienced during launch and ascent.

Launch Vehicle (LV) Environmental Control System (ECS) Impingement

The STIS exterior surfaces shall not suffer damage or degradation when exposed to the LV ECS airflow velocity of 10 m/sec.

6.9 ON-ORBIT DYNAMIC ENVIRONMENT

The STIS shall be designed to handle all permutation of the following linear and angular acceleration requirements simultaneously.

Dynamic Linear Acceleration

The STIS shall survive on-orbit when subjected to the maximum linear acceleration due to nominal thruster firing of ±0.060 m/sec2 in all axes.

Dynamic Angular Acceleration

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The STIS shall survive the maximum angular acceleration of ±2.5E-5 rad/sec2 about all axes.

6.10 THERMAL REQUIREMENTS

The STIS instrument sensor units shall be thermally conductively coupled to the spacecraft for thermal control (TBP).

The STIS instrument heat transferred to the spacecraft shall be limited to 4 Watts [TBR].

The STIS shall meet all of its performance requirements after exposure to air temperature between +5 and +30 degrees C and relative humidity between 30% and 70%.

The STIS thermal design shall employ thermal coatings properties validated to be accurate for materials and mission flight parameters over the lifecycle of the mission.

Flight Interface Design Temperature Limits

When powered “OFF”, the STIS shall be capable of surviving indefinitely when the spacecraft interface temperatures are within the survival limits shown in Table 6-5 without damage or permanent performance degradation.

The STIS shall meet all performance requirements when within the Operational and

Protoflight/Qualification limits shown in Table 6-5.

The STIS shall demonstrate turn on at the Minimum Survival and Maximum operational limits shown in Table 6-5.

Table 6-5 Temperature Limits at Instrument Mounting Interface

Minimum Temperature (ºC) Maximum Temperature (ºC) Operational (In Spec) -10 +40

Protoflight/Qualification (In Spec) -20 +50

Survival (Unpowered) -25 +55

6.11 CHARGED PARTICLE RADIATION REQUIREMENTS

Units containing electronic parts will be exposed to a natural space radiation environment that consists of: (1) trapped particles which include electrons, protons, and heavier ions; (2) particles from solar events (coronal mass ejections and flares); and (3) galactic cosmic ray (GCR) particles.

Definitions

Total Ionizing Dose (TID) - the mean energy deposited by ionizing radiation in a device region divided by the mass of the region. This is often given in units of rad(Si), where 1 rad(Si) = 100 erg deposited per gram of silicon.

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Enhanced Low Dose Rate Sensitivity (ELDRS) - used to refer to a part that shows enhanced radiation-induced damage at dose rates below 50 rad(Si)/s. The enhancement is the result of true dose rate effects.

Non-Ionizing Energy Loss (NIEL) - a measure of the rate of energy…